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Full text of "Influence of vertical-tail design and direction of propeller rotation on trim characteristics of a twin-engine-airplane model with one engine inoperative"

N(\C(^L'l1' ^ 



AER No. L5AI3 



NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS 



WARTIME REPORT 



ORIGINALLy ISSUED 

February 19'4-5 as 
Advance Restricted Report L5AI3 



THE IKFLUENCE OF VERTICAL -TAIL DESIGN AND DIRECTION 

OF PROPELLER ROTATION ON TRIM CHARACTERISTICS 

OF A TWIN -ENGINE -AIRPLANE MODEL WITH ONE 

ENGINE INOPERATIVE 

By Marvin Pitkin, John W. Draper, and Charles V. Bennett 



Langley Memorial Aeronautical Laboratory 
Langley Field, Va, 




WASHINGTON 

NACA WARTIME REPORTS are reprints of papers originally issued to provide rapid distribution of 
advance research results to an authorized group requiring them for the war effort. They were pre- 
** » ' I*' viously held under a security status but are now unclassified. Some of these reports were not tech- 
nically edited. All have been reproduced without change in order to expedite general distribution. 

L - 191 



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VT 



NACA ARR Ko. LfA13 

NATIONAL ADVI30IIY C0?>5?^ITTEE FOR AERONAUTICS 



3XOm79^{ 



ADVANCE RESTRICTED REPORT 

THE Ii??LUINCS G? VERTICAL-TA^L DESIGN AND DlR^CTTO:^ 
0? PROIELLLR RCTATTO:! CN TRIM CHARACTIRI3TICS 
OF A TWIN-ENGIKE- AIRPLANE MOD'sL WITF ONE 
ENGINE INOPERAIiyZ 
By Mai'vin 13 tkiii, Joljn '/■;. Draper, and Charles V. Bennett 

3NMMARY 



Tefrts have been hiade in the Langley free-fllfrht tunnel 
to determine the influence of rriode of propeller rotation 
and vertical-tail design upon the trim characteristic? of 
a model of a tvvin-engirie airplane vtith one engine inoper- 
ative. The test model was mounted on a trim stand, v/hich 
allowed freedom in roll and yaw under conditions sim^ulat- 
ing thore required by the NAGA and the Army Air Force? for 
asymmetric-pcwer operation in flif'^ht. The seven vertical- 
tail desiavis tested included three tails of low aspect 
ratio and of different area, one tm'in tail of low aspect 
ratio, two tails of high aspect ratio and with different 
rudder areas, and one all-mo vaole tail of high aspect 
ratio equipped -.vith a linked tab. All tests vvere made 
v/ith the flaps dov^n. 

The tests showed chat the effect of m:od(. of propeller 
rotation upon tne directional trim characteristics of the 
model operating with a='yrTmetric power was considerable. 
Propeller rotation in v.'hich che upper tips rotate cut- 
board toward the vfing tip (outboard rotation) generally 
created m.ore -severe out-of-trim ccnditicns than ir^board 
rot action. 

The all-movable tJiil design was found to be more 
effective than the other designs tested in nullifying the 
effects of aryrrjnetric pover. The conventional tail de- 
signs with high aspect ratio were mors effective than the 
designs with low aspect ratio in this respect. The single 
vertical tails were generally more effective in tr'mming 
the yawing FiOments created by asy.:itrietric power than 
twin vertical tails of the same asoect ratio and eoual 



KAOA ARR Ko. L5A13- 



area, par-txcul3rl7 when the rudder was free. At p'Tr.all 
anf;le3 of sideslip, however, the mo?Tients caured b^* 
a?yirmetric pov/er Vi/ere more readily tr.'.r'ned by deflecting 
the rudders of the twin tail'^ than by deflecting the 
rudder of a single tail. 

The trlrrmlng effectiveness of the vertical tail in- 
creased al:noet directly V;'lth vertical-tail area but in- 
creased at a decreasin£^ rate with rudder deflection and 
chord. 

When the rudder was free, the addition of dorsal- 
and vsntral-fin areas permitted increases in the asym- 
rnetric power balanced by the tail surface at moderate 
angles of sideslip. 



INTRODUCTION 



of one or more engines of xnultieiigine 
airplanes introduce? a sudden and severe demand upon the 
directional stability o.nd control of those airplanes. 
Such failures result In the instantaneous application of 
large yawing mornents that must be reutralizied either oir 
the rudder control or by the directional stability of the 
airplane. In addition, asyranetric power conditions create 
rc'lling n:3ments that :nust be balanced by aileron deflection 
in order to inaintaln straight flight. This aileron de- 
flection creates additional yawing moments that r'^cuire 
further trimming by the vertical tail surfaces. For multi- 
engine airplanes, then, the asymmetric power condition 
generally im.poses the most severe requlrem.ents for di- 
rectional stabilit:/ and control and to a large extent 
dictates the design of the vertical tall surfaces of these 
airplanes . 

An invef:^tigation has therefore been carried out in 
the Langley free-flight tunnel to provide data concerning 
the relative merits of seven vertical-tail designs and 
two modes of propeller rotation mider conditions of as:/m- 
metric power. ■ The NACA and Army fl-ying-qualitles require- 
ments (references 1 and 2) for directional stability and 
control of airplanes operating with asymm;etrlc power were 
u'^fed to establish the test conditions. The results of the 
invcstigati:!n are reported hrrein. 



NACA ARR No. L5A13 



A J^- scale model of a conventional twin-engine air- 
20 
plane in the r:ed.ixini-bomber class (the North American 

B-2S airplane) was used in the tests. The model was 
mounted on a test stand, which allowed freedom, in yavif and 
roll. The effects of asymmetric power could thus be visu- 
ally observed from changes in the model attitude. 

The seven vertical-tail designs studied in this in- 
vestii^ation varied in either aspect ratio, total tail 
area, rudder area, or general arrangement. Tests vvere 
made with rudders fixed and free, and the effects of 
adding various dorsal and ventral fins v;ere studied with 
the rudders free. The eff-^ct of mode of rotation of the 
operating propeller upon the vertical-tail characteristics 
was investigated for all tail arrangements. All tests 
were made with th^ fla-os down. 



SYMBOLS 



C 



T lift coefficient flijll] 

.-, . _ . ^r.. . . . Rollin 



\ ^ -w / / \ 
-, T -, . , ^r.. . u f Rollins moment \ 
C7 rolling-moment coefficient I ttc^, 

'' ~ /■ _ ^ q-w"w /' 

Cn yawing -mom.ent coefficient (' ^'sving moment \ 
Cv^ rate of change of yav/ing-mom.ent coefficient 



-P 



with angle of sideslip I - - — 




D propeller diameter, feet 

p Gensit;7 of air, slug per c^.'^hic foot 

V free-stream airspeed, feet per second 

Vj. , ,.„ vel:)CitY at end of take-off run, feet t,er 
take-oif ' , ' 

second 

V -^ stalling speed with flacs dovm, feet per second 
m_n _ -c .. ^ X 

q free-stream, dynamic pressure, pounds per square 
foot 



(¥"') 



w 



4 IkACk ARR Fo. L5A13 

T^ effective thrust of one engine, pound r^ 

T) propeller efficiency, percent 

f-;ros3 v;eifTht, pounds 

3y_, v'ing ai-ea, square feet 

bhp brake horsepower of full-scale airplane simulated 
by rrodel 

thp thrust horsepower 

63, corrblr.ed aileron cfeflectlon, do.crrees 

6p I'udder deflection, positive when trailing edge is 
to left, degrees 

5|. flap deflection, positive vvhen trailing edge is 
do^A/n, degrees 

5 elevator deflect:^ on, degrees 
e ^ 

5m tab deflection of all-movable tail, positive 
when trailing edge is to left, degrees 

1, tall incidence of all-T.ovable tall with respect 
to center line of fuselage, degrees 

a angle of attack, degrees 

a+. local angle of attack of x'-ertica] tail, degi'ees 

p angle of siaeslip, degrees 

^ 2/ \ 
A asoect ratio of v^'rtical tail (t-^ /^^ 

V / / 



St area of vertical tail, square feet 

S-j-, balance area o'^ rudder, percent ruddci- area 

S rudder area, square feet 

bf Epan rf vertical tail, feet 

b^^v wing span, feet 



NACA ARR No. L5A13 



APPARATUS 
Wind 'Tannel 



The tests were r.ade in the Langley free-flight tunnel, 
a complete description of vhich will he found in refer- 
ence 5. The tunnel was locked at an angle of pitch of 0° 
for all tests. 



Trim Stand 

All tests were made on a trim stand, v/hich was 
securely festezied to the floor of the wind tunnel. This 
stand was so constructed as to allow the model freedom 
in roll and yaw about the stability axes of the m.odel. 
The stability axes are a system of axes in which the 
Z-axis is in the plane of symmetry of the airplane perpen- 
dicular to the relative wind. The X-axis is in the plane 
of symm.etry perpendicular to the Z-axis. The Y-axis is 
perpendicular to the plane of sjTrm.etry. The origin of 
the stability axes is at the center of gravity of the air- 
plane, which for the present tests vvas located on the 
fuselage center line 25 percent of the mean aerodynamic 
chord behind the leading edge. 

Photographs of the model mounted on the trim stand 
are given as figure 1 and the construction of the stand 
is Illustrated in figure 2. Figure 2 shows that bearing A 
permits freedom in roll and bearing B permits freedomi in 
yaw. A calibrated coil spring was inserted in bearing A 
to provide stability in roll. This alteration m.ade pos- 
sible the measurem.ent of unbalanced rolling moments as a 
function of the angle of bank and thereby facilitated the 
trimjning of these m.oments by means of aileron deflection. 
Both bearings A and B were equipped '."'i th ball bearings to 
keep frictional effects to a minim.um. 

The trimming fin shown in figure 2 -r&s added to the 
trim stand to neutralize the drag yawing moments caused 
when the wind was on by the forvi'ard struts at an angle of 
yaw. Since this fin area was such that the trim stand 
was in complete equilibri'.mi of yawing momorts [Cj^ = C| 

ever the yaw range tested, the trim stand did not affect 
the directional stability characteristics of the m;odel. 



NACA ARR No. L5A13 



Model 



The model used in the investigation was a _!_- scale 

20 
nodel of the North American 6-28 airplane. A three-view 

drawing and a photoe-raph of the riodel are given as figures 

3 and 4, respectively. 

The model was equipped v.ith 2 four-blade propellers 
having a diameter of 8.£0 inches and set at an angle of 
pitch of 20°. Pov;er was furnished hj a direct-current 
controllaiDle-speed electric ir.otor rated, l/s horsepower 
at 15,000 rpm. The left propeller, which was kept in- 
operative during the tests, was so niounted as to windmill 
freely. The right propeller, which v;as used as the oper- 
ating propeller for all tests, was geared to the motor at 
a ratio of 1:5. Provision was made for reversal of the 
direction of propeller rotation. The model v;as equipped 
with partial-span slotted flaps (fig. 3), which were de- 
flected 45° for all tests. 

Sketches of the vertical-tall designs used in the 
Investigation are shov/n in figure 5 and sketches of the 
dorsal- and ventral-fin areas utilized in the rudder- 
free tests, in figure 6. Tail 2 represents the original 
vertical tail surface of the full-scale airplane and is 
considered typical of conventional vertical-tail design. 
The dimensional characteristics of this tail were varied 
to form the other vertical-tail designs. All vertical 
tails were constructed of the NACA 0012 section. In 
order to maintain similitude of hinge -moment character- 
istics as far as practicable, all rudders were of identical 
blunt-nose balance type with a balance area 12.2 percent of 
the rudder area. This type of rudder is of negative float- 
ing tendency and trails with the wind v/hen free. 

The dimensional characteristics of the full-scale 
airplane are given in the following table: 

Vv'ing: 

Area, sq ft 675.90 

Span, ft 72.61 

Aspect ratio 7.80 

Root chord, in '. . - 161.13 

Tip chord, in 67.00 

Mean aerodynam.ic chord, in 120.09 

Root section NACA 23C17 



NACA ARR No. L5A13 



Tip secti or. NACA 4409R 

Percent chora line with zero sv/sepback 33 

Sweepback at leading- edge, deg 4.2 

Dihedral angle , deg 2 

Incidence, deg « 3 

Georr.etric twirt (washout) , deg 2.5 

Taper ratio 2.4:1 

Fuselage J 

Length, ft 54.5 

Section Circular 

Frontal area, sq ft 36.5 

Horizontal tail: 

Total area, sq ft 183.20 

Span, ft 26.35 

Aspect ratio 3. 94 

Dihe dral angle , deg . 

Stabilizer setting, deg 1.50 

Length frorn hinge of elevator to center of 

gravity of airplane, ft 28.90 

Elevator balance area, sq ft 10.63 

Elevator area behind center line of hinge, sq ft 53.00 

Vertical tail 2: 

Total area, sq ft , 74.90 

Span , ft 10 . 68 

Aspect ratio ] . 54 

Length from hinge line of rudder to center of 

gravi ty of airplane , ft 27 . 40 

Pin area, sq ft 35.66 

Rudder area, sq ft 39.24 

Rudder-balance area, sq :t T:' . 14 

Rudder area behind hinge line, sq ft 50.10 

(Pertinent data for tar'ls 1, 3, 4, 5, 6, and 7 are given 
in fig. 5.) 

Aileron (one of tvjo) : 

Area behind hinge line, sq ft ,,....»..... 20.91 

Soan, ft 11.41 

Mean chord, in ....,..,,.. 17,0 



8 NAG A ARH No. L5a13 



Flap: 

Total flap area, sq ft 30.5 

Total span, ft 58.4 

Type Slotted 



SPECIFICATIONS AND CRITZRICNS 

The NACA and Army flight: requirements for multiengine 
airplanes operating with asyrmnetric power were chosen to 
establish the proper test conditions. No separate attempt 
was made to reproduce the Navy specifications for asym- 
metric power because of the close similarity between the 
Navy and the NACA specifications. 



Specifications for Directional Control 

• ■ (Rudder Fixed) 

The NACA and Army specifications (references 1 and 2, 
respectively) for directional control of airplanes operating 
with asjmmetric power are as follows: 

NACA requirement (II-Z) 3 .- "The rudder control should 
be sufficiently powerful to provide equilibrium of yawing 
moments at ^ero sideslip at all speeds above 110 percent 
of the minimum take-off speed 'under the following conditions: 

a. Airplanes with two or three engines: 
With any one engine inoperative 
(propeller in low pitch) and the 
other engine or engines developing 
full rated power," 

Arm.y requirement E-2c ( 1) (c) »•- "The rudder control 
shall be pov.-erful enough to triia a multi-engine airplane 
for straight_ flight with less than 10 degrees of sideslip 
at 1.2 Vq jVo = stalling speed of the airplane, throttles 

"h [_ ""h —i 

closed, gear down, flaps in best tske-off condition] when 
the throttle on an outboard engine is abruptly closed 
(propeller in low pitch) and the other engine or engines 
are developing full take-off power. The flaps shall be 
in the take-off setting, and the gear shall be down...." 



NACii ARR :^o. L5A13 



Specifications icr Directional Stability 
(R VI. cider Free) 



The NACA specification ( requirenient (II-?) 4 of refer- 
ence 1) relating to the rrouiremenbs for directional sta- 
bility with rudder free under asyiumetric power conditions 
is as follows: 

"The yawing moment due to sideslip (rudder free with 
airplane trlmned for straight flight on symmetric power) 
should be .such that straight flight can be maintained by 
sideslippjng at every speed above 140 percent of the mini- 
mum speed with rudder free with extreme asymmetry of power 
possible by the loss of one engine.'' 



Criterion for Vertical-Tail Effectiveness 

".mder As^rmmetric Power Conditions 

Each of tlie specifications previously listed requires 
the dlrection;vl control or the directional stability of 
the sirplsne in question to be sufficiently povi/erful to 
balance the yawing m.oments created by as^Tnmetric power 
under certain specified flight conditions. It follov/s 
that the vertical-r.ai 1 effectiveness in flight m.ay be 
gaged by the maximum amount of as'i^.rrm-aetrlc power which such 
a tail can balance under the specified conditions. In this 
invfc stifration, ther-efore, the miaxlmum asymmetric power 
perrrdssible under the airspeed and trim conditions speci- 
fied by the Army and the NACA. v/as used to evaluate the 
effectiveness of the vertical tails tested. 

It should be observed that the flight specifications 
require that straight flight ">t complete equilibrium of 
lateral forces and momeats be maintained. In order to 
maintain such eauilibrium In flignt, the ailerons miust be 
deflected so thot the rolling moments caused by asymm.etric 
powtr ar-; balanced and the airplane assumes an attitude 
of bank, which nullifies the side force created by rudder 
deflect", on and/or angle of sideslio. Inasm.uch as an atti- 
tude of bank dots not affect the trim req\iirement s of the 
vex^tical tail surface, no attempt '.vas m^ade in the tests to 
simulate the balance of side force by angle of bank. 
Aileron deflection, howevr-r, directly afftcts dii^ectional 
trim by virtue of the yawing mom.ents created by such de- 
flections.. Gonsequeritly , the ailerons wore so adjusted 



10 NACA ARR No. L5A13 



for a] 1 tests as to naintain complete balance of aero- 
dy-nanlc rolling moirents and thereby to simulate flifrht 
cond-itj'ons correctly. 



TESTS 
Test Conditions 



"^he test lift coefficient was established from con- 
sideration of the specified airspeeds in bhe Amy and NACA 
requireirents . These values were converted to 11ft- 
coefflcient forms as follov/s: 

The NACA requirer.ent (II-S) c a (rudder fixed) speci- 
fies an airspeed equal to 1.10 times the take-off speed. 
If '^''t glee -off -^ assumed equal to l'2V,jj^j^, the airspeed 
requirement for this specification is equal to ■^•^^^min* 
If the maximum lift coefficient of the E-SS airplane is 
assumed equal to 2.0, the specified lift coefficient cor- 
responding to 1.22V . Is defined by the expression 

mm 

which equals 1,15. In a similar m.anner, 

the lift coefficient necessary to satisfy IIAGA requirement 
(II-P) 4 (rudder free) was found to be 1.02. The lift 
coefficient necessary to satisfy the Army requirement 
(rudder fixed) was calculated as 1,39. Because it was as- 
sumjed that slight changes in lift coefficient would not 
affect the model test results if the correct values of 
thrust coefficient were used, all tests were r'.m at a con- 
stant ancle of attack of 5°, which corresponded to a lift 
coefficient of 1.10. 

All tests v/ere run at a test velocity of 40 feet per 
second, which corresponds to a tert Reynolds number of 
128,000 based on the mean aerodynamic chord of 0.503 foot. 
The aileron deflections for all tests were adjusted to 
provide equilibrium of rolling moments. 



Test Procedures 

Rudder fixed .- In the tests with rudder fixed, the 
model was mountea on the scand with the rudder deflected 
in the direction that counteracted the yaw caused by 




NACA ARR No. I5A13 11 



apymjrietric power. ^yeasureiuents v^ere then taken of the maxi- 
m\xr, a:r;ount of asyrrinetric thrust the rudder would balance at 
an.'xles of yaw of 0*^ and 10° for rudder deflections of 0°, 
5^7 10'^, 20°, and 30°. 

Rudder free .- The tests with ru:";der free were made by 
measuring the ar.ount of a.=';/mmetric ti.rust and anf-le of yaw 
produced by asynmetric power for various angles of yaw up 
to the an.o:le at which directional instability was encountered 
Tests with rudder free wez'e made of the model with each of 
the following vertical-tail arrangements: 

(1) Vertical tail alone 

(2) Vertical tail olu.s dorsal fin a 
(Z) Vertical tail plus dorsal fin b 

(4) Vertical tall plur ventral fin a 

(5) Vertical tail plus vsntral fin a 

plus dorsal fin a 

The absolute dorsal- and ventral-fin areas required for 
each test were determined from the percentages of the 
vertical tails being tested given in figure 6. No tests 
were made to determine the influence of aixciliary fin area 
upon the characteristics of tw-in tail 4. 

P ovv-er cal c ulations .- Tl-e thrust coefficients that were 
obtained in the tests ci the m.odel were converted to the 
simulated as^>Trjnetric brake horsepovcer of t;:e full-scale 
airplane by m.eans of the rel8.tionship 

vv, thp 
bhp = -^ 

TgV 



or /.-\5/~' 



K^ 



-w 



2 

bhp - -^^-;c-: (1) 



Tel 



5f:0r]p 



12 NACA ARR No. L5A13 



The full-scale propeller efficiency r^ 'va? assun-ed 
to -?e equal to 0.75 for the calculations. Values of wing 
loading '■■'V'/s,,. and propeller diameter D v;ere obtained 

from the full-ccale character! stica of the North American 
E-28 airplane and were equal to 47.5 pound? per square 
foot and 14.7 feet, respectively. The value of the mass 
density of air p was chosen as 0.00258, which is its 
value at sea level mider standard abmospheric conditions. 
SifDsti tution of these values in equation (1) yields the 
relate onship 

T 

thp = 9900 (2) 

Cl" '' 

The values of Cj in equation (2) are those coi^respond- 
ing to tne airspeed specified in the Army and the NACA 
requlreinents and were deteririned as shown in the section 
entitled "Test Conditions." Substituting these values of 
lift coefficient in equation (2) yields the expressions 
defining the conversion of model thrust coefficient T^ 

to the estimated full-scale brake ?aorsepov,-er, which are: 
For n^dder fixed. 



KACA requirement 

Army requirement 

For rudder free, 

FACA requirement 



bhp = 30eOTc (3) 



bhp = 6C70Tc (4) 



bhp = 962CTc (5) 

RESULTS AND DISCUSSION 



The aata obtained in the investigation are plotted 
in figures 7 to 18. Figure 7 sho-z/s the rolling -moment 
coefficients produced by the ailerons used in the tests. 
Figures 3 to 10 present the values of the asyminetric- 
thrust coefficient balanced by means of rudder deflection. 



NACA kRR No. 15^13 13 



Figures 11 to 13 ['ive the vilv.e.s of the asymr.etric-thrust 
coefficient balanced by the yawed model v^'ith rudder free. 
Data showinr the influence of dorsal- and ventral-fin areas 
upon triin characteristics are presented in figures 14 to 13, 

The test data in figures S to 13 vvere rearranged and 
converted to values of full-scale brake horsepower in 
figures 19 to 24. An index to all figures 3S presented 
as table I, 



Effect of Mode of Propeller Rotation 

The mode of propeller rotation in which the upper 
blade tips r.ove toward the fuselage is henceforth designated 
inboard rotation. The rotation in which the upper blade 
tips move out tov>ard the wing tip is designated outboard 
rotation. Almost all conventional airplanes are equipped 
with right-hand propellers. On inultlengine airplanes, 
the direction of propeller rotation with respect to the 
wine tips (inboard or outboard) is therefore determined 
by the location of the propeller. If the right engine 
fails, the direction of the operating propeller rotation 
is inboard and the airplane yaws in a positive sense. 
For left-engine failures, the operating propeller rotates 
outboard and the airplane yaw is negative. The results 
of the present investigation show that use of different 
modes of propeller rotation caused considerable difference 
:' n trim characteristics of an airplane operating under 
asymm.etric power. 

With onl.y one exception, the data presented in 
figures 8 to 13 indicate that the use of outboard propeller 
rotation decreased the values of permissible asyirirretric- 
thrust coefficient balanced by any given vertical-tail con- 
figuration and that this mode of rotation would therefore 
determine the minimum vertical-tail size. The exception 
occurred when twin tail 4 operated under the Army specifi- 
cations (fig. 9); in these tests inboard rotation was less 
favorable than outboard rotation. 

The difference in asyminetric power balanced by a given 
tail arrangement with either of the two modes of rotation 
appeared to increase in m.agnitude with the amount of di- 
rectional stability and of control being applied. The 
largest differences occurred at large rudder angles and for 
tails 6, 6, and 7, which have high aspect ratios. Particu- 
larly large effects of propeller rotation were observed when 
the rudder was free. 



14 NACA ARR No. L5A13 



The magnitude of effect of reversing the propeller 
rotation ha? been illustrated In figure 19. Thi? figure 
presents the calculated values of periaissible brake horse- 
power for both modes of propeller rotation for the repre- 
sentative rudder deflection^ of 20° (figs. lG(a) and 19(b)) 
and for that angle of sides] ip at which directional In- 
stability was encountered in the tests with rudder free 
(fig. 3S(c)). This angle of sideslip was between 10"^ and 
12° for almost all the conditions tested. The results 
presented in figure 19 shov/ that the difference in as;;nn.- 
metric power balanced by the vertical tail for Inboard 
and outboard rotation was about 4C0 horsepower for most 
conditions and was as large as ICOO horsepower for some. 

The effects of changing the direction of propeller 
rotation appear to be ex'plained by the data of reference 
4. Refer-Ence 4 concludes that use of inboard propeller 
rotation with the flaps down caused the slipstream to con- 
verge toward the tail and thereby increased the contri- 
bution of the tail surfaces to directional stability for 
small to moderately large angles of yav'. This slipstream 
displacem.ent would result in a beneficial effect of in- 
board rotation upon the trimming action of the vertical 
tail surfaces, particularly for twin tail 4, which under 
TACA specifications (3 = 0°) appears to be partly im- 
mersed in the slipstream jet. Reference 4 also concludes 
that outboard rotation causes the slipstream- jet to di- 
verge. Consequently, this m.ode of rotation increases 
the tail effectiveness at large angles of yaw but is less 
satisfactory in this respect than the inboard mode of 
rotation for other angles of yaw. This reasoning explains 
the favorable effect of outboard propeller rotation upon 
twin tail 4 when operating at an angle of sideslip of 10°. 
At this angle, ov:ing to its original lateral displacement, 
this tail lies within the slipstream. 



The data obtained in the tests indicate that for 
twin-crginc airplanes equipped with single vertical tails 
and cor.ventional right-hand propellers, the failure of a 
left engine will imtpose the miore severe flight conditions. 
For airplanes eouipped with twin vertical tails, hov/ever, 
the failure of a right engine should prove more critical 
to the fulfillment of the Army requiremients. Similarly, 
it may be reasoned that use of propellers rotating in- 
board on both wings (symmetric rotation) would be advpn- 
tageous for airplanes equipped v/ith single fins both to 
improve tail effectiveness and to m.ake the handling of 
controls similar regardless of tire location of the 



NACA ARR Mo. L5A13 15 



inoperative engine. Conversely, syirjnetric outboard rotation 
should be favorable for airplanes equipped with tv.-in fins. 

Effect of Vertical-Tall De?ign 

Effect of vertioal-tail area .- The effect of varying 
vertical-tail area was obtained frorr a study of the test 
data for geometrically similar tails 1, 2., and 3. The 
data for these tails with rudder fixed vifere converted to 
values of full-scale brake horsepower and plotted in 
figure 20. 

The data of figure £C(a) shov/ that increasing the 
vertical-tail area resulted in increases in the asynunetrlc 
power balanced by a given rudcer deflection at zero angle 
of sideslip, The.'^e increases, however, are not directly 
proportional to the increase in tail (rudder) area, as 
might normally be expected; this lack of proportionality 
Indicates the preoence of secondary slipstream effects 
upon the vertical tail surfaces. Such secondary effects 
are probably produced by the sidev/ash angles generated 
at the tail by inflow into the slipstream jet as well as 
by the more direct effects of slipstream velocity. 
Further Investigation, however, is required to establish 
a complete explanation of these secondary slipstream- 
effects. 

The data in figure 2G(b) illustrate the favorable 
effect upon the asymmetric power characteristics of in- 
creasing the vertical- tail area at an angle of sideslip 
of IC'^ , These data show that, v\?hen the airplane is side- 
slipping, the directional stability of the vertical tail 
surfaces reinforces the action of the rudder control in 
nullifying the effects of asymmietric power, and higher 
values of as^Tnm.etric thrust can therefore be balanced 
by a given vertical-tail arrangement. The magnitude 
of the effects of cirectional stability can be obtained 
from, a study of the curve for a rudder deflection of 
O'^ (fig. 20(b)), which is directly indicative of the 
rudder-fixed directional stability. These data show 
that the directional stabilit:/ contributed by tail 1 barely 
balanced the unstable yawing nomiSnts created by the yawed 
fuselage-wing comoination, ?toking the tail area larger 
than that of tail 1 increased the djrectional stability, 
as v.ould be expected. 



16 • NACA ARR IJo . L5A13 



The e'f'fect?? of iricreas: ng tril area noted "in the 
tests math rudder fixed were also observed in the tests 
Vvith rudder free. Fipure 11 illustrates the influence of 
tail area upon the rudaer-free trim characteristics of 
the model operating under asymmetric power. In this 
figure, the data inJ-icate that freeing the rudder of 
tail 1 was cestabi li.'^ing, as would normally be expected 
since the rudder type employed had a negative floating 
ratio. Because of tiie slender m.argin of stability 
associated with tail 1, the destabilizing action of 
freeing the rudder was sufficient to cause directional 
instability. Making the tail area greater than that of 
tail 1 increased the directional stability contributed by 
the tail surfaces sufficiently to overcome the destabiliz- 
ing effects of the fuselage and, consequently, perm^itted 
increases in the asjrmmetric thrust balanced by the vertical 
tail surfaces. 

£o mp ari son of twi n^ tail and s in gle tail . - T v; i n tall 4 
may be directly compared '.vith tail 2 inasmuch as both 
tails ver? of the same aspect ratio and equal area. Be- 
cause the twin trll was located almiost directly in the 
slipstream, the twin tail was more effective than the 
single tail at zero and small angles of sideslip. Figure 21 
sho.vs that the influence of po^ver at p = 0^' made 
tail 4 almost as effective as tail 3, a surface of equal 
aspect ratj.0 but possessing 50 percent greater area. At 
angles of sideslip greater than 0°, however, tail 4 was 
less effective than tail 2 with the rudder fixed at an 
angle of sideslip of 10^ and with the rudder free (fig. 22), 
These data confirm, trends noted in the past and indicate 
that the directional stability contributed by a twin 
vertical tail la less than that contributed by a single 
tail of the sam.e aspect ratio ana equal area. The in- 
creased directional stability achieved by use of the 
sin.srle tall is partly ascribed to the favorable end-plate 
effect of the horizontal surfaces upon the load charac- 
teristics of ■ the vertical surfaces. In addition, the 
single tail has but one ro^t juncture compared wi th two 
for the twin tail and therefore is less affected by inter- 
ference effects. 

It should be noted that the curves for tail 4 for 
rudder free (fig. 22(b)) do not pass through the origin 
but fall above and belov; It depending on this m^ode of 
propeller rotation employed. These curves indicate that 
reversing the propeller rotation altered the sldewash 
caused by the propeller sufficiently to reverse the local 
angle of attack of tail 4 at small angles of sideslip. 



NACA ARR No. L5A13 



The rer-ults of the te.'^^ts Inuicate that choice between 
single and twin vertical tails would, aepend largely upon 
the pilot's handling of the controls following a sudden 
engine failure. If the rudder c~'ntrol can he applied he- 
fore the airplane reaches a moderately large angle of 
sideslip, the twin-tail den'ign should be more suitable; 
otherw-'ise, the single vertical-tail design would be 
preferable . 

Effect of increasing a-^pect rat ic- The effect of 
increasing aspect ratio was deterrriinedi from a comparison 
of the data obtained with tail 6, a surface of twice the 
aspect ratio of tail 2, with corresponding data for tails 
2 and 3, These data are shown in figure 23 and indicate 
that doubling the aspect ratio of tail 2 has approximately 
the same effect as increasing the area by 50 percent at 
the same aspect ratio (tall 3). This effect is in close 
agreemer.t with the wind-tunnel force data of reference 5, 
which shov;; that doubling che aspect ratio of a surface 
from. 1.5 to 3.0 increased the lift-curve slope from 2.2 
to 3.1. For a given rudder configuration, such a change 
in lift-curve slope v/ould result in an increase in tc^tal 
tail load, or trimming effectiveness, equivalent to that 
obtainable by appro.xim.ately a 50~percent increase in area, 

C om car i_s o n of c onventional t ai 1 and all-m.ovable tail 
wi th linked tabT - The question has been raised whether the 
efficient action of the all-miovable tall reported in refer- 
ence 6 arose largely from the "all-movable" feature or from 
the fact that the tall was of high aspect ratio and had 
the inherent advantages associated with tails of that type. 
For the present investigation, therefore, tests of the all- 
m.ovable tall 5 were supplemented with tests of tail 7, 
which is identical with tail 5 except that tail 7 is of 
conventional - that is, fixed-fin - design. 

A comparison of the effects of tails 5 and 7 upon 
the characteristics of the airplane operating with asym- 
metric power is shown in figure 24. These data indicate 
that the all-movable tell is m.arkedly more effective than 
the conventional tail at zero sideslip with the rudder 
fixed and with the rudder free. At 10° sideslip and with 
rudder fixed, however, the all -movable tail was only 
sllehtly more effective than the conventions 1 tall 
(fig. 24(a)) . 

These test resiAlts miay be explained by use of the 
curves showing typical tall loads (fig. 25) . These curves 



18 NACA ARR No. L5A13 



?how the vai-iatior. of tai?. load vith vertical-tail inci- 
dence and rudder deflection for e. conventional and an 
all-novable tail. The tab area of the all-mo'/able tail 
is assumed equal to the rudder area of the conventional 
tail. The variation of the load with deflection of the 
all-movable tail is indicated by the dashed line in 
figure 25. Thi? variation Is due to the linkage betveen 
the tab and the movable forv;ard surface. The slope of 
the load curve is determined from the linkage ratio 
6iji/i+; which, for the case investigated, was equal to 1,12, 
If the effect of powei' is ignored, the angle of attack 
(tail incidence) of the c r^nventicnal tall at ^i = 0^ is 
also zero. The rudder deflection therefore produces 
changes in load along a path coincidental with the zero 
ordinate. For e:>:a:;riplc , a rudder deflection of 1C° produces 
the tail load corresponding to the load indicated by 
point a, For the all-movable tail, however, a rudder 
deflection of 10° causes a simultaneous change in tail 
angle of attack and tab deflection and produces the load 
indicated by point b. Consequently, at zero sideslip, 
the all-movable tail is capable of producing much larger 
yawing moments v-ith wl : ch to balance the effect of asym- 
metric power than the conventional tail. 



At moderate angles of side-^lip (1C° to 15"-^), the con- 
ventional tail operates in the high-lift region of the 
lift curve of the ta-* 1 and consequently produces tail 
]oads of an order comparable with those px'-oduced by the 
all-movable tail. The conventional tail may conceivably 
produce tail loads even greater than those of the all- 
movable tail because the conventional tail r's unrestricted 
in the use of rudder. The all-movable tail, hovifcver, is 
limited for a given linkage ratio to the rudder deflection 
that produces the tail Incidence at maximum, lift. Further 
deflection would cause the entire surface to stall. 

In balancing the effects of esymnptric pov;er, the 
superiority of the all-mjovable tail to the conventional 
tail was most msrked in the rudder-free tests. This su- 
periority can be ascribed to the fact that the hinge- 
moment characteristics of the all-movable tail force the 
entire tail tD float against the wind v/hen free (positive 
floating ratio) and crnsequentl^r Increase the directional 
stability of the airplane. In considering the advantages 
of the ail-r.iovable vertical tail over the conventional 
tail, it should be observed that "snaking" oscillations 
may be induced by control-surface friction with improperly 
designed tails having positive floating ratios. (See 
reference 6.) 



NACA ARR No. L5A13 19 



decreasing 



rudder chord is shown by the te:^t data for tails 6 and 7 
(figs. 10 and 1?). These data show that, although the 
rudder of tail 7 had only one -half the area and one -half 
the cliord of the rudder of tail 6, the rudder of tail 7 
balanced approxirrately two-thirds as much asymmetric power 
at zero sideslip and approximately seven-eighths as much 
power at 10*^ sideslip as the rudder of tail 6. These data 
are in agreement with conventional trends because it is 
known that decreasing the rudder cnord increases the yaw- 
ing mcm.ent per unit rudder area, 

With rudder free, tall 7 balanced a greater amount 
of asyiTiiretric povv^er than tail 6, which indicated a favorable 
effect of reduced rudder area upon the rudder-free direction- 
al stability. This action occurred because the rudders of 
tails 6 and 7 are of the type that trail with the v.'ind and 
so reduce the directional stability when set free. Con- 
sequently, tail 7, because of its smaller rudder area, 
created smaller destabilizing mom.ents when the rudder was 
set free and so balanced a greater amount of as^nnnetric 
power. 

Effect of rudder deflection ,- The data obtained in 
the tests showed that increasing the rudder deflection 
increased the amount of asymimetric power balanced by the 
vertical tails at a decreasing rate. 

Effect of dorsal and ventra l fins.- The data illus- 
trating the 'effect of adding dorsal- and ventral-fin areas 
to tails 2, 3, S, €, and 7 are presented in figures 14 to 
18. No data are presented for tail 1 because the addition 
of dorsal and ventral fins did not noticeably lessen the 
directional instability associated with t'lvis tail arrange- 
ment. 

The test data indicated that the addition of auxiliary 
fin area Increased the directional stability at large 
angles of yaw and thereby increased the miaxim-um. amount of 
asymmetric thrust balanced by tne tail surfaces vi/hen the 
rudders were free. Increases in maxir.iuiii asymanetric thrust 
of the order of 30 to ICO percent were observed in the 
tests. 

The addition of ventral-fin area was generally found 
to be m.ore effective than the addition cf an equal amount 
of dorsal-fin area. The use of a com.blnation of dorsal- 
and ventral-fin areas (dorsal, a and ventral a) -was 



20 NACA ARR No. L5A13 



pen^rplly rore effec"i"ive than a sin.'^le dorral f;.n of the 
same total area (dorral b) . 



A"jleron deflections 



e ction s req ui red to t rim agymmetr l£ 
entativo plot of total aileron deflec- 



thru'-t . - A repi'esL.-. ^^^„ ^^ „,„„^ ...^ ... ..„,.^.> 

tions required to trjm the rolling raorrents created hj 
asymmetric thrurt is presented In figure 26. These 
deflections were alvi'ajrs obtained bv equal up-and-dov-Ti 
movoF.ents of the ailerons. Calculated values ai-e also 
presented in figure 26. These calculations were made 
by using the method presented in reference 7. The calcu- 
]ated lift increments created b^/ the operating propeller 
were rrultiplied by the lateral arm of "the propeller to 
obtain rolling moments, which v.'ers converted to aileron 
deflections required to ti-iim by use of the datr. in figure 7, 

The results presented in figure 26 shov; that, although 
the scatter way considerable, the test data agreed fairly 
well with the calculated values and Indicated that moder- 
ately large aileron deflections woula be required to main- 
tain straight flight undei- as:\Tiimetric power conditions. 



OOFCLnSTONS 



The following conclusion"^ were drawn from trim tests 
of a twin-engine-airplane m.odel operating under asymmetric 
power (single-engine) conditions specified by the NACA and 
Army Ai r Forces: 

1. The direction of rotation of the operating pro- 
peller had an important effect upon the asymmetric power 
that could be balanced by a given vertical- tail design. 
Single vertical tails were mo5t effective when the oper- 
ating propeller was rotating inboard, Tv;in tails, however, 
were most effective when the operating propeller v/as 
rotating outboard. 

2. An all-movable vertical tail of aspect ratio 3 
with, a linked tab was more effective than the conventional 
tail of the same aspect ratio and equal area in balancing 
asymmetric pov;er, particularly when the rudders were free. 
The all-movable tall was markedly superior to the con- 
ventional vertical tail -^f normal aspect ratio (1.5). 

3. The single vertical- tail desiQ-ns generally balanced 
a greater am.ount of asyrnme trie power than twin vertical 



NACA aRR xMo. L5A13 



tails of the rair.e aspect rat5_o and equal area, particularly 
when the rudder war; free. At sinall angles of sideslip, 
however, it was, possible to balance more power by rudder 
deflection of the twin tails than by rudder deflection 
of a single tail. 

4. Increasing the aspect ratio of a vertical tail 
resulted in increasing its triirining effectiveness under 
asymmetric power conditions by an amount proportional to 
the accompanying increase in lift-curve slope. 

5. The tri'Ti.-.iing effectiveness of the vertical tail 
surface increased almost linearly with the vertical-tail 
area. Increaslrg .'■udder deflection and rudder chord in- 
creased the trirrjT'l::r effectiveness of the vertical tail 
under asymmetric concitlons at a decreasing rate. 

6. When the ^udder was free, addition of dorsal- 
and ventral-fin areas increased the capacity of the 
vertical tail surfaces to balance asymmetric-power effects 
at moderate angles of sideslip. 



Langley Memorial Aeronautical Laboratory 

National Advisory Committee for Aeronautics 
Lanslev Field, Va. 



22 NACA ARR No. L5A15 



REPER^NCDS 



1. '.7j.].ruth, P. R.: Requirements for Satisfactory Flying 

dualities of Airplanes, NACA ACR, April 1941 / 
(Clasfification changed go Restricted Oct. 1943.) 

2. Anon,: Stability and Control Requireirents for Air- 

planes. A/iF Specification No. C-1815, Aap. ?1, 1943. 

3. Shortal, Joseph A., and Cstc-rliout, Clayton J.: 

Prelirrinary Sta^^ility and Goncro] Tests in the 
NACA Free-Plight V^nd Tunnel and Correlation with 
Full-scale Fll^lit Tests, FACA TK No. 810, 1941. 

4. Pitkin, t.'arv'.n: Free-Flieht-Tunnel Investigation of 

the Effect of Mode of Prooeller Rotation ' upon the 
lateral-Stability Characteristics of a Twin-Engine 
Airplane Model \vith Sinale Vertical Tails of Dif- 
ferent Size. NACA ARR No. 3 JIB, 1943. 

5. Zir.'!ir.erman, C. H.: Characteristics of Clark Y Airfoils 

of Small Aspect Ratios. NACA Rep. No. 431, 1952. 

6. Jones, Robert T., and Kleckner, Harold F. : Theory and 

ireliriinary Flight Tests of an All-I^ovable Vertical 
Tail Surface. NACA ARR, Jan. 1943. 

7. Pass, H. R.: V.'ind-Tunnel Study of the Zffects of 

Propeller Operation and Flap Deflection on the 
Pitching Moments and Elevator Hinge ivlotnents of a 
Single -Enshrine Parsuit-Tyoe Airplane. NACA ARR, 
July 1942 ; 



NACA ARR No. L5A13 



23 



TABLE I.- INDEX TO FIGDRES 



Figure 


Inscription 


Remarks 


^ 


Photographs of test model mounted on trim stand In Lfingley free- 
fllffct tvinnel 


Model with tall 2 


2 


Sketch of test model mounted on trim stand, which permitted 
freedom in yaw and roll. In Langley free-flight tunnel 




3 


Three-vlee drawing o£ w».-5cale twin-engine icodel teaCed In Langley 
free-flight tunnel with asymmetric power 




« 


Photograph of twln-englnc model used in trim teste In Langley free- 
flight tunnel 


Model with tall 2 


S 


1 — 

Plan-form and dlnenalonal characteristics of seven vertical tails 
tested on a ^-scale model of a twin-engine alrulane In the 

Langley free-flight tunnel 




6 


Various fin arrangements tested with vertical tails on a ^-scale 

model of a twin-engine airplane In the Langley free-flight 
tunnel 




Figure 


Te«t apeel- 
tleatlont 


Taet 
condition 


Tall 
arrangement 


Operating-propeller 
rotation^ 


Cnrre 


Remarks 


Tall 


Dorsal [Ventral 


7 






2 


j 


Propeller off 


C[ against 6^ 


Aileron calibration 










8 


a 


NACA 
(p = 0°) 


Rudder 
fixed 


1 to 
3 






Inboard and outboard 


Tg against 5^ 


Elrectlonal-control 
run 






b 


NACA 

(p = 10°) 


Rudder 
fined 


1 to 
3 








Tj against 6^ 


Do. 








9 


a 


NACA 
(p = 0°) 


Rudder 
fUed 


4 








Tg against 6p 


Do. 








b 


Army 
(p = 10°) 


Rudder 
fixed 


4 








T, against 6, 


Do. 








10 


a 


NACA 

(0 = 0°) 


Rudder 
fined 


S to 

7 








7g against 6^ 


Do. 










Army 
(p = 10°) 


Rudder 
fined 


S to 

7 






(i-j_ ..__■ 


T^ against 6^. 


Do. 










11 




NACA 
(Rudder freej 


Rudder 
free 


1 to 
3 






Outboard 


Tj against p 


Dlrectlonal- 
stablllty run 








NACA 
(Rudder free) 


Rudder 
free 


1 to 
3 






Inboard 


Tf against p 


Do. 






12 




NACA 
(Rudder free) 


Rudder 
free 


4 






Outboard 


T, against p 


Do. 








NACA 
(Rudder fre^ 


Rudder 
free 


4 






Inboard 


T5 against p 


Do. 






13 




NACA 

(Rudder fre^ 


Rudder 
free 


5 to 
7 






Outboard 


Tc against p 


Do. 








NACA 
(Rudder fre^ 


Rudder 
free 


5 to 
7 






Inboard 


Tg against p 


Do. 






14 




NACA 
(Rudder fre^ 


Rudder 
free 


2 


combinations^ 


Outboard 


Te against p 


Effect of dorsal- 
and ventral-fin 
area 




NACA 
(Rudder free) 


Rudder 
free 


2 




Inboard 


Te against p 


Do. 


15 




MACA 
(Rudder free) 


Rudder 
free 


3 


^= 


Outboard 


Te against p 


Do. 




NACA 

(Rudder fred 


Rudder 
free 


J 


do 


Inboard 


Tj against p 


Do. 



iRlght propeller operative. 

^f7nm*Tlnit1 "n* tested ape tell alone, dorsal 



a, dorsal and ventral e, Testl*! «. *orsal b. 



.NATIONAL ADVISORY 
COimiTTEB FOR AERONAUTICS 



NACA ARR No. L5A13 



24 



TABLE I. - INDEX TO FIGURES - Concluded 



Figure 



Tejt speci- 
fications 



16 



NACA 
(Rudder free) 



Teat 
condition 



Rudder 
free 



Tall 
arrangement 



Tall ! Dorsal [Ventral 



Tall aloB*, 
dorsal a, 
dorsal a and 
ventral » 



Operating-propeller 
rotation^ 



Outboard 



Curve 



Tq against p 



Effect of dorsal- 
and ventral-fin 
area 



Remarks 



NACA 
(Rudder fre^ 



Rudder 
free 



All 
combinations^ 



Inboard 



Tg against p 



Do. 



17 



NACA 

(Rudder free) 



Rudder 
free 



do 



Outboard 



T(. against p 



Do. 



NACA 
(Rudder free) 



Rudder 
free 



do- 



Inboard 



Tf against p 



Do. 



18 



NACA 
(Rudder free) 



Rudder 
free 



do 



Outboard 



Tj against p 



Do. 



NACA 
(Rudder free) 



Rudder 
free 






Inboard 



Tq against p 



Do. 



19 



NACA 
O = 0°) 



Rudder 
fixed 



1 to 
7 



Inboard and outboard 



bhp for 

6p = 20= 



Effect of mode of 
rotation 



Army 
(p = 13°) 



Rudder 
fixed 



1 to 
7 



-do- 



bhp for 
6- = 20° 



Do. 



NACA 
(Rudder free) 



Rudder 
free 



1 to 

7 



-do- 



bhp for 
10<kp<120 



Asymmetric power 
balanced at verge 
of directional 
divergence 



20 



NACA 
(P = 0°) 



Rudder 
fixed 



1 to 



Outboard 



bhp against 
tall area 



Curves of various 
rudder deflections 



Army 
(3 = 10°) 



Rudder 
fixed 



1 to 
3 



-do- 



-do- 



Do. 



21 



NACA 
(3 = 0°) 



Rudder 
fixed 



2 to 
4 



Inboard 



bbp against C^ Comparison single 
and twin tall 



NACA 

(P = 0°) 



Rudder 
fixed 



2 to 
4 



Outboard 



bbp against Oj 



Do. 



22 



Army 
(3 = 10°) 



Rudder 
fixed 



2. 4 



Inboard and outboard 



bhp against 6^ 



Do. 



NACA 
(Rudder free) 



Rudder 
free 



2, 4 



-do- 



bhp against p 



Do. 



NACA 
(9 = 0°) 



Rudder 
fixed 



2,3, 

6 



Inboard 



bhp against Cj, 



Effect of aspect 
ratio 



NACA 
(Rudder free) 



Rudder 
free 



2, 3, 

6 



-do- 



bhp against p 



Do. 



24 



NACA (p»0°) 
Army (J=10°) 



Rudder 
fixed 



5, 7 



Outboard 



bhp against Q, 



Comparison of all- 
movable and con- 
ventional tall 



25 



26 



NACA 
(Rudder fr««J 



Rudder 
free 



5. 7 



-do- 



bhp against p 



Do. 



Tall load 
against i^ 



Illustrative of 
principle of all- 
movable tail 



Inboard and outboard 



Tg against &a 



Aileron deflections 
required to trim 



Ifilght propeller operative. 
^Combinations tested are tall 



ilone, dorsal a, dorsal and ventral a, ventral a, dorsal b. 



NATIONAL ADVISORY 
COmtXTTKE FOR AERONAUTICS 



NACA ARR No. L5A13 



Fig. 1 





Figure 1.- Test model mounted on trim stand in Langley 

free-flight tunnel. 



NACA ARR No. L5A13 



Fig. 



z! o 




o 
a 

iij 



V) 

< 

1. 



Ill 
a 
o 



lij 

3 



I 



UI 




rr 




h 


-I 




_i 


V 


o 


III 


a: 


_1 




o 


n 


.ii 


7' 


< 


< 


_1 




7' 


5 




< 


n 


> 


z 




< 




1- 




(/) 




S 




a: 





Q 

H 
Z 

O 

5 



LU 
Q 
O 



cn 



I 

LU 



i 



NACA ARR No. L5A13 



Fig. 3 



1734" 




Propeller diam.j 8.^" 



Wire landing gear 



Figure 3. - Three-view drawing of l/zo -sca/e twin- 
engine model as ijsstecc //? Long/ey -fnse- 
f//g/if- Tunnd iv/^Ji asym/r^ei-r/c poiver. 



NACA ARR No. L5A13 



Fig. 4 




NACA LM^L 27625 




NACA IJ*U. 27627 



v.. 



Figure 4.- Twin-engine model of B-2.6 airplane used in trirr 
tests in Langley free-flight tunnel. 



NACA ARR No. L5A13 



Fig. 5 




^95 



Cut-out for 

horizontal 

tail 




Arn^^psrr.^^ \ 



550 



S^ ffiu x pj ji . ^ 



Tail 



122 



5^^ y^/^^n^ 5>6.d^0l^ 3 



11.0 



12'Z 



5U 15^ 



J_ 



16.5 



12.2 



I.5U 



11.0 



12.2 



T5U 



^eo 



1 1.0 



12.2 



^ 



II.Q 
12.2 



3.0 



25d^0 2$.0 



JIO 

12.2 



3.0 




(one of two) 

ItLTab 



o. Tail 5 pivots about 0.27 MAC 
li of GUI taib /6.^3" behind e.g. 

NATIONAL ADVISORY 
iw COMMITIEt FOR AERONAUTICS 




lQi\5 (a/l-movab^) 



Tail 6 



Tail 7 



Figure 5. -Plan -form and dimensional charactehetics of seven 
vertical tails tested on a 1/20- scale model of a 
twin-engine airplane /n tha Langley free -f//(^ht 
tunnel. 



NACA ARR No. L5A13 



Fig. 6 




NACA ARR No. L5A13 



Fig. 7 
































i; 

I— <■ 






'\ 




















? 




\ 


\, 


















3 






\ 
























M 


\, 
























\ 


\, 
























\ 


K 
























\ 


\ 
























\ 


V 
























\ 


\, 
























\ 


^ 
























\ 


^ 
























\ 


























\l 



8 



O 



s 



Q 



CM 

o 



<3 



O 



'9 



q O 

'iUdioijidOD lUddiouz-duiiiOf^ 



^ 



03 8^ 



y) c: 



I 



NACA ARR No. L5A13 



Fig. 8a, b 



o Operating pmpeJ/er rofat/nQ ouftxxuzi 

H -Operating propeller roTof/ng /ntjoara 




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flight rurtter defject/onj 6rj cleg 
(b) Army 5pea-fical/on5 ^=/o). 
F/guns 3.- Asymme^r/c-potyer chanychrf^Hcs of a iw/n- 
enp/ne - airplane moPel eQ.u/ppGcL m^ yer//ca/-fa// 
designs I , Z, and 3. cL suc/i fhaf f/ie ro/t/rxp momsnf^ 
equal Oj df^dS") <^-0°j o(^S° left prope//er ^/nd- 
millingj rudder fixed.. national advisory 

COMMITTEE FOR AERONAUTICS 



NACA ARR No. L5A13 



Fig. 9a, b 



O- 



Operaf/ng propeller rofaf/ng oultnord 
Operating propeller roratinq inboard 




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to ^o 30 



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f?/ght rudder deflect ion ^ dr, deg 
(b)Army 5peaf/caf/on5(/3-/0^). 

figure 9- Asynnmetric-px^/ver choracfensf/cs c/a tw/n-engine- 
airplone model eouipped vMiTh vertical- fail design -4. 
6c 3ucfi fhafftie rolling moments equal o ,6^45°. 6e-'0: 
oc--j°-/eff propeller tA/irdrniiling , rudder fixed. 

NATIONAL AllVisoSy 
"— . •;n ni.i.(KilAuflcS 



NACA ARR No. L5A13 



Fig. 10a, b 



O operotinq propeller rotal/ng outboord 

+ Operating propeller rotating inboard 



Hinge line 

OflQU 




^HinQelim 
,oftab 



Jail 5 




Toil 6 





lai I deflection L,om , , ^, , 

dr/L.^i.i2 ' 17ighi rudder deflection, 6 r cleg 
(b)Army specif/cafiom (.S=/OV' 
Figure /o - A^ynimctr/c -power characteristics or a fw/n-engine- 

airplane model eguipped witH verticol-tail designs S, 6, 

and 7. 6q ^uch tnoff/^e rolling moment5 equa/Oi 6f = 4S] 6^ = O'- 

oc- 5". le// propeller windmil/ing j roa'cfpr p/xecT. 



NACA ARR No. L5A13 



Fig. lla,b 



-^Mode/ airecr/onOJ// un^/otDt 




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(b) Operating p/vpe/ter m/at/ng /r}tx)ard. 
T/gure /t .-/^s/m/m/r/c -pother ct/orocter/st/cs ot a /mn-e/ig/ne- 

Qjrpanc mode/ egw/uped with i/ert/cat- /q// cfes/g/7S /,2,Qnd 

3. 6a sudi that the ro////ig nxmenfs equatO;S^--4S°Se=0''- 

cc--5l/ett propeller windmilling- rudder Tree. 



NACA ARR No. L5A13 



Fig. 12a, b 



O 



^ Model d/rect/ 0/1011/ unstoD/e 
— Vertical fojl ojone 



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rooe of- fiu>oy 



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NATIONAL ADVISOHY 
COMMIIlEt FUR AERONAUTICS 



-to , Q ^ , lo^ ^n JO 
Angle of sides iip.p, aeq 



(b) Cperating ,oropeller m/ar/ng //?^oara. 

f/gore /Z.- As/mmefr/c-porver choracten^^f/c^ of a tt^in-engine- 
Qirolonp model eguippecl lA/itti vertical -tail design ^^. . 
6o so'c/7 r/xjt r/7c rot ting moments eguotO;<S^--'^S;'Se-0;CK^ 5; 
t eft propeller windmilling ^ rudder tree. 



NACA ARR No. L5A13 



Fig. 13a, b 



^Moctel c//recf/onQJ/y unstc^/e 



line of 

IQ/I 




H/nge 
-//ne 
of lad 



lQll5 




Tail 6 




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siJ^ICS 




Ang/e of ^/desl/pj (3 > deg 
(b) Qperal/ng propeller rofaf/ng inixxjrcL . 
Figure 13 .- /Isymmelric -pother chamd'er/5f/cs cf a ft^/n-eng/ne- 
ojrplane model egu/pped i/v/f/i yerfjcoJ-tojl di^/gns ^,6,aria7. 
cfa sucf) ifnf //le ro///ng momen/s equal 0;Sf--45°6s--0lcc^5° 
leff propeller mnOm/lling ■ ruMer free . 



NACA ARE No. L5A13. 



Fig. 14a, b 



^Mode/ a/rectionoJl/ unsToJDie 




TaiJZ 



Dorm/d 
a 



DoooJ VenM 



— a a 

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(t>) Operaf/ng prQOG//er m^a^/ng /n/xxjroL. 

F/gure /^ . -£/^/ecf o/ ab/:soJ-Qrct vpntroJ-f/n oreos upon 
t/ie Qsymme^nc-poyver choractenst/cz of a fw/n-engine- 
ourpiane model egu/pped yyifh vert/ ca/- feu I c/es^gn 2. 
6a.^uch that the /xDJ/ing momentz equal O; 8f = 45; 6^-0] 
ex -- 5°; Jeff prcpe/fer windmill ing ; ruddet free. 

NATIONAL ADVISORY 
COMMITTEE FOR AERONAUTICI 



NACA ARR No. L5A13 



Fig. 15a, b 



Mode/ c/>r£c//onQJ/y ans/alj/e 

-raz/B 




Dorsal ti 
a 



o- 



Oorsa/ \/en/raJ 



8— -3 a 

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/a cO ^ Ja , -^O , JD 60 NATIONAL ADVISORY 

Anq/e of sjdeslipj^^dsg cummittee for aeronautics 

(b) Opera/jng prope/fer rok3///x/ /nboanX . 

F/gure 15 . - Pf/ecf of dorsaJ- onol venr/aJ-fJn areas upon the 
Qsy/vmefnc-poyyer cnorac/ef/^f/cs of a. fwjn-^ng/ne.-curp/a/)e 
model equjpped w/rh veP/cal-fojl design J. da Such thof /t>e 
roj/ing nnomenfs equo/ O; 5^=45^, 6e-0] cc= 5° /eft propeller 
i/^tndmil/ing, rudder free. 



NACA APR No. L5A13 



Fig. 16a, b 



■Model directionaJly unstoJble 




ro//5 



o- 



Dorsil YenfraJ 

- o. — 

- a a 



Ventral Q 



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Angle cf fia^Sj/0^(3,&/ Wvmittee for aeronautics 
r2)j Ope far /ng propeller rokji/ng /ndosrct. 

^iqure 16 . - Effecf of dorsal- and venfral-f/n areas upon //)e 
Qsymmefnc - poller choracfensHcs of a fmn - eng /ne- 
at rpiane model eqa/pped w//h \/srfical-i-ci/l des/gn J~. 
da such lf)af //le rolling momenis equal 0;6f-.45° 6^-0] 
cx-S°j left propeller mndm/II/ng) rudder free . 



NACA ARR No. L5A13 



Fig. 17a, b 



Mcx:lel cf/recf/anoJ// u/xsfad^ 



rQ//c> 




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Dorsal \^nffnl 



— Ol cl 

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TU /(U CXJ JU </u ou 

F/gure n.'Effect Of abrxd-OfiCJ ventral -f/n a^eos u/uon tne 
asymmefnc -/x>i^r chorocter/st/c^ of aiw/n-enginc-ojrp/one 
nrxJel eguz/Queo/ \Nitt} v^rficoJ-tcul design 6. c5a such ttiat 
the filing mo/mehh equal O-, Sf = 45° de = 0°,c!: = 5 J /eft propeller 
wrCrDiWirg ^ rudder f/^. 









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NACA ARR No. L5A13 



Fig. 18a, b 



Mode/ c//rect/cnQj/y unsfo/Dt 




ro//r 



o- 



Dorsal ventm.1 



V- 



— a — 

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iO , 2Q 3Q, <0 , SO 
'Oes/// 



(D) Opera fjng pmpe//er rofaf/ng joboord. 
ngur&l8.-£'ffoof o/aorWil'Qrxi venrraJ-f/n arsas upon the asymme/rjc- 
pO)/ver c/iarac/enstic5 of a fyi//o-eng/ne.-Qirplane rncdel eQu/ppsa 
With verlicoJ-rajl dssjgn 7. 6^ 3uch r/nf fhe ro/Z/ng moments equal O; 
6f^s°; c5e-o°- <^-^''; /e/f propel /or wind/7v/lJr)gpudoJer free- 



NACA ARR No. L5A13 



Fig. 19a, b.c 




Hinge 
/jnc-, 



^lOil 4- \Ta\\ 5 TOJI 6 Tail 7 

pie c/ two) laJhmoiaUe) 



3eco- 



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Inbcyard /'dof/on 

U7X Out^tyoard rokj/ton 

_ — £5ti mated by 
extrofiolaf/on of 
c/aJa 



(a) MCA specifications (dr^^ZO"). 



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NATIONAL ADVISORY 
COMMITTEE FOR AERONAUTICS 



t 



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' / 2 J 4 5 6^ f ^-jQJi 

(c)A/ACA :5pecificaf/ons ■ rudder free (JO'^^-^/i'^- 

Figure /9. 'Fffed of mode of prcpeJIer rotaf/on upon fhe permiss/b/e 
asymmetric dra/re ttorsepoiver., oaJcu/ated from /est cloJa for 
the North Americon B-2& curptane cperatinq on one engine 
under N/\CA ond Artny flighf specifications. 



NACA ARR No. L5A13 



Fig. 20a, b 



H/nge 
line 



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NATIONAL ADVISORY 
COMMITTEE FOR AERONAUTICS 



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Vert/col- toil area, >St , fract/on /[/jng Q/ea. 
(b)Arm\/ 5p6c/f/ca//on5 0S=/O°J. 




model Test Oai^cLTor the Norf/? Afnencan B'2Q airptone 
op&xmfig on one engine under t\/AC/l and /\rmy ft/ght 
^pecificol/on5 Operat/ng p/tpe/ter nofat/ng ou/txord . 



NACA ARR No. L5A13 



Fig. 21a, b 



2^00 




4 8 IZ J^ ZO Z4 2^ 

Right rudder cfef/eci/on , di-, dep 

f/gumll -Companion of t/ye asymmefnc-poouQr c/xfrac/er/^- 
f/C6 ofana/rp/ane egalpped uj///? fu/^n one?' ^//^ig/e 
\/ert/cQ/1a//-5./VAC^.^oec/f/co/-/oo5j /S-O". 



NACA ARR No. L5A13 



Fig. 22a, b 




/nboa/Tii rofoHon 

I I I I I 

Qu^doard. /v/a^/on 



Ouidoarct rofQf/on_ 



'Inboa/n rotation 



lONn AEVISOHY 



COIIMIT EE (OR /ERONAUTICS 



I 
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I 

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Right rudder d.efJection, c^ , cLeg 
(a) /^r/ny specifications Qs ~/0°). 



I200 



QOO 




400 



a /z 
Ang/G of sicLes/ip, ji3 , cLeg 
(t) /VACA specifications ; rudder free. 



Figure 22. - Comparison of the asynimetr/c -potver 
characteristics of an airplane equipped with 
twin and sing/e vertical tails and at various 
angles of sideslip. 



NACA ARR No. L5A13 



Fig. 23a, b 






Q 
Q 

V 
O 

i 



MOO 
2000 

I6C0 

1200 
WO 
4O0 



O 



Tail Aspect ra+/o 
Z /■S4 




I 



4 Q 12 16 20 24 23 32 
P/ghf- rudder def/ecHon j cfr > deg 
(a)A/ACA 5pGC/f/ca//or)5 f^^O% 




Ang/e of sides l/p^ jff j deo 
(b) A/AC A 6pec/f/ca-f/on5jru<daer -free. 

Figure 23. - Comparison of ihe asymmetric —power 
characferisHcs of an airplane sQuipped wifJi 
high- and loiv-aspeci--raiio \rcriicQl iaib. Inboard 
rohiiion. 



NACA ARR No. L5A13 



Fig. 24a, b 



IdOCf 




4 8 IZ l<o lO Z4 Z8 3z JS~ 

/?/g/7f rodder der/ecT/on, d'^,(^^o'eg 




~1 ? 6 S /o /i A? /6 /c3 
/Iry/g" of sideslip ^ Q ^ deg 

f/gure 24.- Compor/son of //?<? o5)/m/7)e-/-nc -poiVtsr c^arac- 
■fensHcs of a fw/n -eng/ne - a/rp/ons mocLel eQo/ppecL 
wifh on Qll-moyad/e verf/caJ fa/l w/fh /j/iMed /ad 
and with a 3/ngJe concern zona/ fo/L Operaf/ng 
props/ /er faming ouf^oard. 



)i 



NACA ARF No. L5A13 



Fig. 25 



-Lxdi cur\/e ^r oyZ-movcUble ioj\ (cL/c li^) 
■Load, ouj-i/e ^or conu&ntionoJ ^L^^\ x . pQ" 



I 

o 




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^Load. produced, 
by a/I- mos^ojb/e 



Load, produced Jby convvn-fional 
fail af /S^O% cfr = /O' 



IS 



ao 



5 lo 

VerficoJ -fcul incidence., i^ ") 
Ang/e of ^ides/ip ^ ^ cCeg 
Angle o-f oi+ack of fair, ocA 

F/gur&2.5.- Typical lood cun/es for /he oJI-mo^<xbJe 
and com^enfionoJ types of vert/cai frul. 

NATIONAL ADVISORY 
COMMITTEE FOR AERONAUTICS 



i 



NACA ARR No. L5A13 



Fig. 26 






24 



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.Calculated 



NATIONAL ADVISORY 
COMMITTEE FOR AERONAUTICS 



.08 .16 .24 .3£ (K> 

Asymmetric - thrust coefficient^ Tc 



.4S 



Figure Z6~ Aileron deflections reqaired to trim 
rolling moment creoled by asymmelric 
po\A/er (5=0°, operating pnopeller 
tvtatinq outt>oord. 



UNIVERSITY OF FLORIDA 

IIHIII'IIIIIIIIII 



3 1262 08106 463 5 



UNIVERSITY OF FLORIDA 
DOCUMENTS DEPARTMENT 
120 MARSTON SCIENCE LIBRARY 

P.O. BOX 117011 .^,,n<^x 

GAINESVILLE. FL 32611-^011 USA